air captured from the tips of a compressor of a gas turbine engine is diverted from the engine core and can be collected for auxiliary uses. gas path separation can be achieved using part-span shrouded compressor blades or using blade tip cut-outs conforming to an airflow dividing annular shroud. In a preferred application for the present invention, the gas turbine engine is the auxiliary power unit of an aircraft. This permits compressed air generation for Auxiliary power unit oil cooling and for compartment pressurization without loosing significant mass flow to the engine core, while eliminating the need to provide a separate active cooling system such as an engine driven fan.
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1. A gas turbine engine compressor, comprising:
a rotor adapted to rotate about a central axis, the rotor having a hub and rotor blades extending radially from the hub; an annular compressor casing being concentric with said central axis and defining an outer wall; said rotor blades having tips wherein at least part of said tips extend to said outer wall, and said blades having end portions near said tips; said outer wall extending upstream of said rotor, permitting substantially unobstructed and undivided fluid flow communication between an exterior air source and said rotor; a stationary annular shroud within said compressor casing and concentric with said central axis, extending downstream from said rotor, said annular shroud having a leading edge which is substantially parallel to said central axis; a first annular duct defined within said annular shroud; said annular shroud and said outer wall defining a second annular duct; said first duct permitting core fluid flow communication between said rotor and a compressor outlet; and said second duct adapted to supply air from at least said end portions of said blades for auxiliary use.
14. A gas turbine engine compressor comprising:
a rotor adapted to rotate about a central axis, the rotor having a hub and rotor blades extending radially from the hub; an annular compressor casing being concentric with said central axis and defining an outer wall; said rotor blades having tips wherein at least part of said tips are in close proximity with said outer wall, and said blades having end portions near said tips; said outer wall extending upstream of said rotor, permitting substantially unobstructed fluid flow communication between an exterior air source and said rotor; an annular shroud within said compressor casing and concentric with said central axis, extending downstream from said rotor; a first annular duct defined within said annular shroud, said first duct permitting core fluid flow communication between said rotor and a compressor outlet; said annular shroud and said outer wall defining a second annular duct, said second duct adapted to supply air from at least said end portions of said blades for auxiliary use; a centrifugal impeller downstream of said rotor; and a stator interposed between said rotor and said centrifugal impeller.
10. A gas turbine engine compressor comprising:
a rotor adapted to rotate about a central axis, the rotor having a hub and rotor blades extending radially from the hub; an annular compressor casing being concentric with said central axis and defining an outer wall; said rotor blades having tips wherein at least part of said tips extend to said outer wall, and said blades having end portions near said tips; said outer wall extending upstream of said rotor, permitting substantially unobstructed and undivided fluid flow communication between an exterior air source and said rotor; a stationary annular shroud within said compressor casing and concentric with said central axis, extending downstream from said rotor; a first annular duct defined within said annular shroud, said first duct permitting core fluid flow communication between said rotor and a compressor outlet; and said annular shroud and said outer wall defining a second annular duct, said second duct adapted to supply air from at least said end portions of said blades for auxiliary use, wherein said gas turbine engine is an auxiliary power unit, and wherein an auxiliary use comprises providing cooling air for a passive oil cooling system for said auxiliary power unit.
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12. The gas turbine engine compressor as defined in claims 10 wherein said rotor blades have at least a flange extending from both sides of each blade to form part-span shrouds concentric with the central axis and aligned with said annular shroud.
13. The gas turbine engine compressor as defined in
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17. The gas turbine engine compressor as defined in
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20. The gas turbine engine compressor as defined in
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The present invention relates to gas turbine engine compressors, and more particularly to capturing air from a compressor tip for auxiliary uses. More specifically, it pertains to using the captured air supply for the passive cooling of an auxiliary power unit.
Gas turbine engine powerplants are used in the vast majority of aircraft flying today. Most large commercial aircraft include an auxiliary power unit (APU), generally a small gas turbine engine, often mounted in the aft tail section of the aircraft, which provides electrical power and pressurized air for aircraft environmental control systems when the aircraft is on the ground, and is also used to start the main engines of the aircraft. APUs require external cooling and are lubricated by oil that is generally cooled by an air cooled oil heat exchanger.
Active cooling systems are most often employed to provide this cooling air, and are typically comprised of a fan used to push air through the oil cooler and across auxiliary power unit components. These fans are driven at high speeds by the APU through relatively complex shaft and gear assemblies. The mechanical complexity and high operating speeds of these fans increase the possibility of failure of the cooling system, which would eventually lead to APU shutdown. Active fan cooling systems therefore significantly reduce the reliability of an auxiliary power unit, and add considerable cost and weight. While various passive cooling systems exist, they often require ducting air from the exterior of the aircraft, and fail to be able to provide compressed air for other uses.
Various systems used to separate compressor airflow are known. U.S. Pat. No. 5,357,742 issued Oct. 25, 1994 to Miller, for example, discloses metering cooling air exhausted through a turbojet laminar flow nacelle system, to cool the core engine compartment. Air bled from the entry to the core engine compressor drives a turbocompressor pump which draws cooling air through the laminar flow nacelle system and into a manifold surrounding the engine. This system has the disadvantage of requiring a separate pump to provide the compressed cooling air.
Separating airflow from the exit of a centrifugal compressor is also known. In U.S. Pat. No. 2,696,074 issued Jan. 2, 1953 to Dolza, an engine and torque converter cooling system having a two stage impeller and an annular diffuser is disclosed. Air is diverted from the main air stream flow, into either impeller stage. One or both of the impeller stages can be engaged. Two separate diffuser inlet nozzles accept air from each impeller stage and feed two diffuser chambers, one intended to cool the torque converter and the other the engine. The inlet airflow to the impeller is separated from its inlet and is selectively directed to one or both impeller stage inlets.
Passive cooling solutions particularly for auxiliary power units are numerous. U.S. Pat. No. 6,092,360 issued Jul. 25, 2000 to Hoag et al., discloses an APU passive cooling system in which cooling air is drawn into the engine compartment through an opening located in the rear of the aircraft. An eductor mounted before the exhaust duct of the engine, draws compartment air through the oil cooler, which in turn draws atmospheric air in through the aft opening.
Therefore, while methods of auxiliary power unit oil cooling and compartment pressurization exist which eliminate active cooling systems, there is a need for an APU built-in passive cooling system capable of providing compressed air for cooling and other uses. While some attempts have been made to use compressors as a source of cooling air, none employ the engine core compressor for a cooling system that does not require additional ducting of cooling air from the exterior of the aircraft.
It is an object of the present invention to supply cool air from the compressor of a gas turbine engine to be used for a means other than power generation.
It is another object of the present invention to fulfil the cooling and compartment pressurization requirements of an auxiliary power unit in an aircraft.
Therefore, in accordance with the present invention, there is provided a gas turbine engine compressor, comprising: a rotor adapted to rotate about a central axis, the rotor having a hub and rotor blades extending radially from the hub; an annular compressor casing being concentric with said central axis and defining an outer wall; said rotor blades having tips wherein at least part of said tips are in close proximity with said outer wall, and said blades having end portions near said tips; said outer wall extending upstream of said rotor, permitting substantially unobstructed fluid flow communication between an exterior air source and said rotor; an annular shroud within said compressor casing and concentric with said central axis, extending downstream from said rotor; a first annular duct defined within said annular shroud; said annular shroud and said outer wall defining a second annular duct; said first duct permitting core fluid flow communication between said rotor and a compressor outlet; and said second duct adapted to supply air from at least said end portions of said blades for auxiliary use.
Further features and advantages of the present invention will become apparent from the following detailed description, taken in combination with the appended drawings, in which:
The rotor assembly 12 rotates axially about the engine center axis and generally serves to increase the velocity of the incoming air. The rotor 12 is principally comprised of a central rotor hub 17 and a plurality of radially extending rotor blades 16 having tips 25. The stator 14 is comprised of a plurality of axially extending stator vanes 11 which redirect the air flow exiting the rotor blades 16 and increases the static pressure of the air. The gas path 22 is shown for the main compressed air duct to the engine core.
The rotor 12 can be a one piece unit, an "Integrated Bladed Rotor", comprising the central rotor hub 17 and the integral rotor blades 16. Traditionally, however, individual blades 16 are mounted on the central hub 17 using a fir-tree style attachment well know in the art, and can have either shrouded or non-shrouded tips. Throughout the compressor, the gas flow path decreases in cross-sectional area in the direction of flow. This reduces the volume of the air as compression progresses. The centrifugal compressor stage comprises the impeller 27, a single forging often composed of titanium that generally has a plurality of blades 29 and an integral hub 31, and a diffuser 45. The blades 29 guide the axial air toward the outer circumference of the impeller, increasing the velocity of the air by means of the high rotational speed of the impeller. The subsequent diffuser 45 serves to straighten the airflow and to convert the high velocity, high kinetic energy into low velocity, high pressure energy. The use of axial and centrifugal compressors is well know in the art.
In the present invention, a flow dividing annular shroud creates a bifurcation in the compressor exit gas path, providing an alternate externally directed gas path for pressurized cooling air which can be used for purposes other than power generation, such as APU oil cooling and compartment pressurization requirements.
In the first embodiment shown in
For a compressor portion 110 of an alternate embodiment shown in
In one application of the present invention, it is proposed to use air diverted from the tips of the axial or centrifugal stage compressor of an auxiliary power unit, for air cooled oil cooling and compartment pressurization requirements. The present invention would therefore provide a passive cooling system which eliminates the need to provide a separate fan running as an accessory to the engine, and is self-contained within the engine. This translates into a significant cost and weight saving as well as improved product reliability. The auxiliary air produced by the compressor could equally be collected and used for multiple other uses.
The embodiments of the invention described above are intended to be exemplary only. The scope of the invention is therefore intended to be limited solely by the scope of the appended claims.
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